Aeroacoustic Experiments in the Langley Low-Turbulence ...

NASA / TM-2002-211432

Aeroacoustic Experiments in the LangleyLow-Turbulence Pressure Tunnel

Meelan M. Choudhari, David P. Lockard, Michele G. Macaraeg, Bart A. Singer, and

Craig L. Streett

Langley Research Center, Hampton, Virginia

Guy R. Neubert, Robert W. Stoker, and James R. Underbrink

Boeing Commercial Aircraft, Seattle, Washington

Mert E. Berkman, Mehdi R. Khorrami, and Shelly S. Sadowski

High Technology Corporation, Hampton, Virginia

February 2002

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NASA / TM-2002-211432

Aeroacoustic Experiments in the LangleyLow-Turbulence Pressure Tunnel

Meelan M. Choudhari, David P. Lockard, Michele G. Macaraeg, Bart A. Singer, and

Craig L. Streett

Langley Research Center, Hampton, Virginia

Guy R. Neubert, Robert W. Stoker, and James R. Underbrink

Boeing Commercial Aircraft, Seattle, Washington

Mert E. Berkman, Mehdi R. Khorrami, and Shelly S. Sadowski

High Technology Corporation, Hampton, Virginia

National Aeronautics and

Space Administration

Langley Research CenterHampton, Virginia 23681-2199

February 2002

Acknowledgments

Many individuals contributed to various aspects of this work. The authors especially would like to thank the crews atthe Langley Low-Turbulence Pressure Tunnel for making this work possible.

The use of trademarks or names of manufacturers in this report is for accurate reporting and does not constitute an I

official endorsement, either expressed or implied, of such products or manufacturers by the National Aeronautics andSpace Administration.

Available from:

NASA Centerfor AeroSpaceInformation (CASI)7121 Standard Drive

Hanover, MD 21076-1320(301)621-0390

National Technical Information Service (NTIS)

5285 Port Royal Road

Springfield, VA 22161-2171(703) 605-6000

Contents

Nomenclature ...................................................................................................................................................... iv

1. Introduction ..................................................................................................................................................... 1

2. Experimental Details ...................................................................................................................................... 1

2.1. Experimental Facility ............................................................................................................................. 1

2.2. Airframe Model ...................................................................................................................................... 2

2.3. Instrumentation ...................................................................................................................................... 3

2.3.1. Pressure-Sensitive Paint ................................................................................................................ 3

2.3.2. Surface Pressure Ports ................................................................................................................... 3

2.3.3. Unsteady Surface Pressure ............................................................................................................ 3

2.3.4. Phased Microphone Array ............................................................................................................. 4

3. Results ............................................................................................................................................................. 8

3.1. Flap Side Edge ....................................................................................................................................... 8

3.1.1. Mean Flow ..................................................................................................................................... 8

3.1.2. Unsteady Surface Pressures ........................................................................................................ 11

3.1.3. Acoustics ...................................................................................................................................... 14

3.1.4. Flap-Side-Edge Modifications .................................................................................................... 17

3.2. Leading-Edge Slat ................................................................................................................................ 28

3.2.1. Mean Flow ................................................................................................................................... 28

3.2.2. Unsteady Surface Pressures ........................................................................................................ 29

3.2.3. Acoustics ...................................................................................................................................... 30

3.2.4. Slat Modifications ....................................................................................................................... 32

4. Concluding Remarks .................................................................................................................................... 37

References .......................................................................................................................................................... 38

111

Nomenclature

c speed of sound

cruise-wing chord

Cp pressure coefficient

f frequency

M Mach number

Re Reynolds number based on

SPL sound pressure level

x streamwise distance along wing

c_ angle of attack of configuration

6f flap deflection

6 s slat deflection downward

iv

Abstract

A phased microphone array was used in the Langley Low-Turbulence

Pressure Tunnel to obtain the radiated acoustic JieldJbom high-lift wing

conJ_gurations that included a slat and a part-span flap. The data

included noise localization maps and acoustic spectra.

1. Introduction

Airframe-generated noise is an important component of the total noise radiated from commercial

aircraft, especially during the approach. Recent studies by Davy and Remy (ref. 1) on a scale model of an

Airbus aircraft indicate that the high-lift devices and landing gear are the main sources of airframe noise

when the aircraft is configured for approach. Earlier tests on a model of a McDonnell Douglas DC-10

aircraft also identified the high-lift system as an important airframe noise source (ref. 2).

Dobrzynski et al. (ref. 3) performed full-scale experimental studies in an open-jet wind tunnel on a

portion of a wing equipped with a high-lift system. They found that both the leading-edge slat and the

side edge of the trailing flap contributed significantly to airframe noise.

An extensive experimental and computational effort to study the various mechanisms associated with

airframe-generated noise continues at NASA Langley Research Center (ref. 4), including tests performed

in the NASA Langley Quiet Flow Facility (QFF) (refs. 5-7). Many aspects of the work are described

elsewhere (refs.8-10); this report documents the experimental conditions and the salient results.

Section 2 discusses the experimental facility, the model geometries, and the instrumentation used.

Section 3 summarizes the results and important conclusions from the various tests.

2. Experimental Details

This report documents three wind tunnel entries. The first entry, in August 1997, focused on noise

emanating from the flap side edge. This wind tunnel entry is logged as Flap-Edge Noise Test 403. The

second entry, in April 1998, explored some noise-reduction technologies for flap-side-edge noise and then

examined slat noise. The respective portions of the tests are logged as Flap-Edge Noise II Test 409 and

Slat Noise Test 410. The third entry, in June 1999, tested some noise-reduction techniques for slat noise

and then studied landing-gear noise. These studies are logged as Acoustics-Flap Noise Test 420 and

Acoustics-Landing Gear Test 421, respectively. Details of the landing-gear test are not discussed here.

2.1. Experimental Facility

The tests involved a series of aeroacoustic experiments performed in the Langley Low-Turbulence

Pressure Tunnel (LTPT). The LTPT is a closed-loop wind tunnel that can operate at pressures up to

10 atm, thereby providing the capability to test at variable Reynolds numbers without changing the Mach

number. This capability, critical to these tests, enabled clear determination of which effects were

Reynolds number related and which could reasonably be expected to exist in flight. A detailed description

of the tunnel is given by McGhee, Beasley, and Foster (ref. 11). Some details pertinent to these tests aresummarized below.

Unfortunately, the wind tunnel is not well suited for aeroacoustic tests. The tunnel walls are all hard

and the test section is not large; its dimensions are 3 ft wide, 7.5 ft high, and 7.5 ft long

(0.98 m x 2.46 m x 2.46 m). The tunnel has a passive boundary layer control system for venting sidewall

boundarylayersthroughporousendplates,butaeroacousticconcernsledtotheuseofsolidendplates.Thetunneltemperaturewascontrolledbytunnelcoolingvanes,butautomaticcontroloftunnelpressurizationwasdisabledto avoidnoiseintroducedbycontrolvanes.Anotheraeroacousticconcernwasthemodelsupportarcsectorthatextendsfromthetunnelfloor to theceilingatthedownstreamendof thetestsection.Thearcsectorisdesignedto supportinstrumentationfor measuringdetailsof themodelwake.Leftuntreated,openingsin thearcsector(suchasscrewholesandrecessesformountingequipment)wereamajorsourceofnoise.Simplytapingtheopeningswasinsufficient;thehighoperationalpressuresoftenburstthetape.Ultimatelyametalplatewasfabricatedto coverseveralmajoropenings,all remainingopeningsinthearcsectorwerecaulkedwithacryliccaulk.

ThetargetReynoldsnumbersbasedonthemodelmeanchordin thecruiseconfigurationwere3.6,7.2,14.4,and19.2x 106.ThetargetMachnumberswere0.125,0.2,and0.3.Notall entrieswererunatallReynoldsandMachnumbers.BecausestrongReynoldsnumbereffectswereobservedonlyforthelowestReynoldsnumber,mostof therunsusedaReynoldsnumber(Re)of 7.2x 106or greater.TheMachnumber(M) of 0.2wasselectedasthebaselineMachnumberfor thesetests;theotherMachnumberswereusedprimarilytodeterminescalinglaws.

2.2. Airframe Model

The two-dimensional (2D) NASA Energy Efficient Transport (EET) wing described by Morgan

(ref. 6) served as the basic model for the tests. In the cruise configuration, the wing has a chord g of

21.65 in. (55 cm); this is the chord with which lengths will be nondimensionalized. In the 1998 and 1999

tests, the same leading-edge slat was used. This slat has been described by Lin and Dominik (ref. 13). No

leading-edge slat was used in the 1997 test. Two flap configurations were used in the tests. In both the

1997 and 1998 tests, a part-span trailing flap was used. The part-span flap had a span of 19 in. (48.26 cm)

and a 6.5 in. (16.5 cm) chord (30 percent of chord). For the 1999 test, a full-span trailing flap was desired.

Unfortunately, a full-span version of the part-span flap previously tested was not available. Rather than

build an entirely new full-span flap, the full-span flap of Lin and Dominik (ref. 13) was used. This flap

was also a 30-percent chord flap, but its contour differed from that used in the previous tests.

Schematics of the high-lift configurations used in the 1997 and 1998 tests are shown in figures l(a)

and (b), respectively. Gaps, overlaps, and deflection angles were all defined consistent with figure 6 of

Lin and Dominik (ref. 13). The slat and flap overhangs were set using blocks machined to match the main

element contour, whereas the gaps were set by rearranging spacers at the feet of the support brackets.

Modifications to the basic wing will be described briefly in the noise-reduction sections below.

In the tests that focused on flap noise (the 1997 and the first part of the 1998 tests), the model was

mounted in the wind tunnel with the suction side of the wing facing the wind tunnel ceiling. For the other

tests the model was inverted so that the pressure side faced the wind tunnel ceiling. These arrangements

were selected to center the acoustic array (described below) over the region of interest in the particular

tests. For the slat portion (i.e., the second part) of the 1998 test, the model was also moved downstream

approximately 29 in. (73.66 cm) to position the slat more nearly under the acoustic array. The translation

required the construction of a manual turntable and mounting system. For the 1999 tests, the standard

turntable and mounting system was used.

J L-

(a) Configuration for 1997 test and first part of 1998 test.

(b) Configuration for second part of 1998 test.

Figure 1. High-lift wing section. Slotted flap shown with flap deflection 8f 30 °. Slat shown with slat deflection8s 30°.

2.3. Instrumentation

Various types of measurements were taken during these tests. Pressure-sensitive paint was used during

the 1997 entry to determine flap-side-edge vortex paths. This technique was not used during later wind

tunnel entries. Surface pressure measurements were taken for portions of all three entries. Extensive

phased microphone array data were taken during all three entries. Additional aerodynamic data were

taken during aerodynamic portions of the 1997 and 1998 entries and will be reported elsewhere.

2.3.1. Pressure-Sensitive Paint

Pressure-sensitive paint (PSP) was used to help determine vortex trajectories associated with the flap

side edge. The paint changes color in response to the local static pressure. Vortices near a solid surface

leave a low-pressure footprint on the surface. A team from what was then McDonnell Douglas

Corporation applied the paint, took the photographs, processed the data, and plotted the mean pressure

fields on the flap surfaces. Data were obtained from cameras directed to the top of the flap (camera 3) and

the side edge of the flap (camera 4). The processed data on the flap top and side surfaces are in digital

format and can be viewed from a variety of angles. A more detailed analysis of the data is presented in thesection entitled "Mean Flow."

2.3.2. Su_ace Pressure Ports

The models included surface pressure ports for obtaining static pressure at discrete locations on themodels.

2.3.3. Unsteady Su_ace Pressure

Several of the models were outfitted with fluctuating pressure transducers for obtaining unsteady

pressure measurements on the wing elements. Most of the unsteady pressure data acquisition was

performed with a mix of equipment and staff from High Technology Corporation and The Boeing

Company. During the 1997 test, the signals from the unsteady pressure transducers were low-passed

filtered at 40 kHz. During the 1998 test, the unsteady pressure transducer data were reliable to a

frequency of about 70 kHz.

2.3.4. Phased Microphone Array

The microphone array and processing software were supplied by The Boeing Company. Underbrink

and Dougherty (ref. 14) and Dougherty (ref. 15) describe the use of logarithmic spirals in phased

microphone arrays. This spiraled array improves the signal-to-noise ratio available compared to

conventional microphone array configurations. The improved signal-to-noise ratio allows for the

acquisition of acoustic data in hard-wall tunnels. Mosher (ref. 16) and Mosher et al. (ref. 17) address

some additional issues that arise when using arrays in hard-wall wind tunnels.

The microphone array was installed in the ceiling of the wind tunnel test section by covering three

removable ceiling sections with a single 1/2 in. thick plate faired at the upstream and downstream ends.

Microphones were placed in the regions where the ceiling sections were removed. The positions of the

ceiling sections are indicated in figure 2. The large array employed 52 microphones embedded in the

plate with Boeing custom-designed flush-mount microphone adapters. In the 1997 entry, Briiel and Kj_er

(B&K) 4136 microphones were connected to B&K flex-necks to adapt half-inch Larson Davis

preamplifiers (model number 900B) to the quarter-inch microphones. Larson Davis 12-channel power

supplies (model number 2212) were located inside the pressure chamber (but outside of the test section)

and the analog signals were brought out of the pressure chamber by means of connector patch panels that

are part of the tunnel wiring infrastructure. Data provided by B&K indicate that the B&K 4136

microphones experience a sensitivity reduction of 25 dB at 50 kHz and 6 atm. as compared to 1 atm.

Largearray center

[

1.122 in.-_

+y _4 Air flow

+x High-frequencyarray center

(15.8095 in.,-12.75 in.)

/East window

Figure 2. Schematic of plate and removed ceiling sections in LTPT. View is from above test section; array centersare centroids of microphones actually used in corresponding arrays.

SubsequenttestsusedB&K 4136WB1437microphones,whichhavea substantialsensitivityimprovementathighfrequenciesunderpressurizedconditions.Themaximumlossin sensitivitywiththenewmicrophoneswasabout15dBatapproximately40kHz.Theuseof thenewmicrophonesimprovedthedynamicrangeofthemeasurements.The1999testhadanadditionalequipmentupgradewiththeuseofB&Kpreamplifiers(modelnumber2670/WH3188)andBoeing-built16-channelpowersupplies.

In all tests,datawereacquiredastime serieswith HewlettPackardHP-3565dataacquisitionhardware.Thesystemconsistedof anHP-35654Acontrolmodule,anHP-35653Asourcemodule,twoHP-35659ASCSI(smallcomputersysteminterface)diskcontrollermodules,andupto 60HP-35652Binputmoduleswith 102.4kHz databandwidth.Theinputmodulescontainedall thenecessarysignalconditioningfor qualitydigitaldataacquisition,includinganalogantialiasingfilters,16bit A/D (analogto digital)converters(dynamicrangeof 75dB).Theconvertershaverangingcapabilitythatenablesdynamicrangeoptimizationwhendigitizingthedata.UsingBoeingcustomdataacquisitionsoftware,digitaldatawerestreamedin realtimetotwoSCSIdiskdrivesof2 Gbeachandsuccessivelyuploadedtothehostcomputer(anHP-9000/385workstation).Onthehostcomputertheywerewrittentobinaryfilesasdigitized(rawinteger)values,alongwithinformationtoconvertthedatatovolts.

Thedesignof thelarge-aperturearrayusedin all threetestsis bestunderstoodby examinationoffigure 3. An oddnumberof microphonelocations(13)wereequallyspacedarounda seriesof10concentriccircles.Intersectionsof thecircleswitha logarithmicspiral(locationsmarkedwithsolidcircles)werechosensothateachmicrophonelocationwouldoccupyanequalapertureareaonthearraypanel(exceptfortheinnermostcircle,whichwaschosenindependently).A completearraywouldincludemicrophonesat all the indicatedlocations.However,limitedaccessin thewindtunnelrestrictedthemicrophonepositionsto thosethatcomfortablyfit in theremovedceilingsections,whicharerepresentedbythethreerectangularregions.

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20

15

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10

15

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Figure 3. Schematic of microphone array. Open circles show desired locations of microphones; rectangular boxes

show regions in which microphones were actually included; solid circles show intersection of logarithmic spiral withconcentric circles.

Thearraywasinitiallydesignedto workwellupto afrequencyof40kHz,buttestsindicatedthatthestrongestsidelobeat80kHzwasnotsignificantlystrongerthanthatat40kHz.Formostfrequencies,theworstcasesidelobein aplaneparalleltothearray,atadistanceequaltothedistanceof themodelfromthearray,wasapproximately8 dBdownfromthepeak.However,designinganarrayonapanelwhererelativelylargeregionsarenotinstrumentedyieldedanundesirableoutcome,asdescribedbelow.

In thepast,untilandincludingthiswork,thearraydesignprocessinvolvedtheevaluationof arrayresolutionandsidelobecharacteristicsonaplanarsurfaceneartheregionof interestfor thetest.Worstcasesidelobesweredeterminedatmanyfrequenciesonaplanarsurfaceparalleltothearraysurfaceatthesamedistancefromthearrayasthemodelwouldbeduringthetest.Thisprocessdidnotdetectlargeout-of-planesidelobesthatwerelaterdiscoveredfor thisarray.Fortuitously,theout-of-planesidelobeswereabove-and-aftandbelow-and-forwardof asourceneartheflapedge,andthereforedidnotinterferewithflap-side-edgemeasurements.

In designinganarray,anotherimportantissueshouldbeconsidered.Volumetricbeamformingshowsthatisosurfacesof constantnoiselevelareellipsoidswithmajoraxesorientedonalinefromthesourcetothephasecenterofthearray.Thisorientationmeansthattheworstresolution(abilityto separatecloselyspacedsources)is alongthismajoraxis.In casesinwhichmultiplenoisesourcesneedtobeconsidered,careshouldbetakentoensurethatthesourcesarenotcollinearwiththephasecenterof thearray.Suchcollinearitywouldmakedistinguishingthedifferentnoisesourcesverydifficult. For instance,thissituationcouldariseif a landinggearweremountedonthewingandthearraycenter,thewheelsofthelandinggear,andtheslatgapwereallcollinear.

The1999windtunnelexperimentalsoincludeda60-microphonesmall-aperturearray,locatedin theeastwindowplateofthewindtunnel(seefig. 2).Thisarraywasdesignedtoworkwelluptoafrequencyof 80kHz,andthereforeis sometimesreferredtoasthehigh-frequencyarray.Eachconcentriccircleofthearrayincluded11equallyspacedmicrophones.Theirrelativepositionswerechosenby usingaproceduresimilartothatusedfor thelarge-aperturearray.Thepanel'slayoutofmicrophonesisshowninfigure4. Inthefigure,theasterisksmarkthelocationof microphonesassociatedwiththelarge-aperture

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8 6 4 2 0 2 4 6 8 10Streamwise window coordinates

Figure 4. Schematic of microphones mounted in east window plate. Circles show locations of microphones for

small-aperture array; asterisks show locations of microphones used in large-aperture array.

array.Wheretheselarge-aperturemicrophonelocationsweretooclosefor mountingpurposesto thedesiredlocationof a microphonefor thesmall-aperturearray,thesmall-aperturemicrophonewasexcluded.Theconfigurationof thearraywaschosenfromanalyticalandempiricaldataregardingtheslatnoisesource.All microphonesin thesmall-aperturearraywereintendedtobewithinthemainlobeof theslatnoisesource,wherethemainlobewasobtainedfromSinger,Lockard,andBrentner(ref.8).

Calibrationswereperformedacrossthefrequencyrangeof interestwithmultiplespeakersources,includingalow-frequencypipespeaker,asupertweeter,andanionophone.Thespeakerswerecarefullypositionedandthepositionsdocumented.Theknowngeometryenableddeterminingtheas-installedchannel-to-channelphaseandamplitudedifferencesbetweenthe end-to-endmeasurementsystemchannels.A spectrumfromareferencemicrophone(acentralmicrophonein thearray)wascapturedforusein calibratingthearraylevel.Pressure-responsedatawereappliedto thereferencemicrophonetoaccountfor microphonesensitivityvariationsunderpressurizedconditions.In the1999test,a sonicdigitizerwasusedfor thefirst timeto determinetestgeometry.Thismethodprovidedmoreaccuratedeterminationof testgeometrythanthepreviousplumbbobandtapemeasuremethods.However,thetightconfinesoftheLTPTtestsectionmadeusingthesonicdigitizerarraydifficult.

Processedacousticarraydataprovided"sourcelocalizationmaps"foreachfrequencyprocessed.Poorspatialresolutionin thedirectionnormalto thearrayunderminedassociatinganoisesourcewitheithertheupperor lowersideof theflapbasedonlyonasinglefigure.Inpractice,viewingmultipleplanesofsuchdatais usefulandoftengiveshintsthatstronglysuggestonesourceovertheother,butpreciselocalizationin thenormaldirectionis notyetpossible.However,theacousticarrayhadgoodspatialresolutioninplanesparallelto thearray,sothelocalizationmapswereusefulforlocalizingconcentrateduncorrelatedsourcesin thecontourplane.Theacousticarraypostprocessingwasnotsuitedforlocalizingspatiallyextendedsources(likeconventionaltrailing-edgenoise)norsourcesthatwerecorrelated.Foruncorrelatedsources,sidelobestypicallyappearedabout8dBbelowthepeakamplitude.Fordistributedsources,thesidelobescouldappearatamplitudesmuchclosertothepeak.

Moredifficultieswereencounteredwhendataat differentfrequencieswereto becompared.Mostoften,spatiallyintegratedspectrawerechosenfor study. Theintegratedspectrawere1/12-octaveinfrequencyandhadbeenintegratedoverthespatialdomain.Gridpointswithnoiseamplitudesmorethan8dB downfromthepeakwerenot includedin theintegration.Theexclusionof thesepointsfromtheintegrationwasintendedtoavoidintegratingthesidelobecontributions.

Althoughthisprocedureis generallybelievedto provideusefulinformation,onoccasionit hasthepotentialto bedeceiving.In particular,ata givenfrequency,considerabaselinecasein whichnoiseemanatesataboutthesameamplitudefromalargeportionof thespatialdomain.All of thegridpointswill be includedin the integratedresultfor thisbaselinecase.If somemodificationto thebaselineconfigurationweretointroduceanewhigh-amplitudesourcethatwashighlylocalizedin thedomainandmorethan8 dBstrongerthanthebackgroundnoisethatpreviouslyexisted,mostof thegridpointsthatpreviouslycontributedtotheintegratedresultswouldbediscountedin thenewintegrationbecausetheiramplitudewouldbemorethan8 dBdownfromthepeak.Dependinguponthedetails,thisphenomenoncouldresultin theloudernoisecaseactuallyintegratingto a lowervaluethanthebaselinecase.Thissituationwasavoidedinmostcasesbyconsideringbothpeakandintegratedresultsduringdatareductionandevaluation.

In spiteof thesedifficulties,theacousticarraysprovidevaluableinformationfor unravelingthephysicsandevaluatingvariousnoise-reductiontechniques.

3. Results

3.1. Flap Side Edge

Noise generated near the side of the flap is related to the unsteady flow associated with the side-edge vortices.

Therefore, considerable effort was expended in understanding the details of the vortices in the vicinity of the flapside. Both computational and experimental efforts focused towards understanding this flow are reported by

Berkman et al. (ref. 9). Section 3.1.1 describes the mean flow. In section 3.1.2, experimentally obtained surfacepressure information is supplemented with knowledge of the flow obtained from numerical simulations.

Section 3.1.3 describes key features of the acoustic field. Finally, section 3.1.4 concludes with a review of theperformance of noise-reduction techniques for the flap side edge.

3.1.1. Mean Flow

A rendition of the main element and the flap is shown in figure 5. The dashed line is approximately

1.0 in. (2.54 cm) inboard of the flap side edge and indicates the approximate location of the line of

pressure taps used to obtain the pressure distribution over the main element and the flap as illustrated in

figure 6. The second suction peak on the flap is caused by strong suction of the vortex. Pressure

distributions farther from the flap side edge do not show the second peak. The main element used was not

the cruise-configuration main element, but the high-lift configuration without the leading-edge slat.

Essentially continuous pressure distributions from the PSP measurements are shown in figure 7. The

pressure distributions confirm that a dual vortex system exists near the flap side edge. In the current tests,

the stronger vortex forms on the flap top surface. This result differs from observations made with a

different configuration that was tested in the NASA Langley Quiet Flow Facility (QFF) and the

Ames 7- by 10-Foot Subsonic Wind Tunnel (refs. 18-20). As shown in figure 8, five-hole probe

measurements from the QFF show that the stronger vortex in those experiments developed on the flap

side edge. The flap side-edge vortex then merged with the flap top surface vortex (ref. 18). However, the

general trajectories of the vortices are similar. In the current experiments, the low-pressure region near

the side edge of the flap top surface in figure 7 reveals that the flap-top vortex moves slightly inboard as it

moves downstream. The weaker vortex forms on the flap side edge close to the bottom comer where the

boundary layer on the flap separates. It grows in strength in the downstream direction and eventually

migrates onto the flap-top surface, where it merges with the flap-top vortex.

The delayed growth of the flap-side-edge vortex relative to that in the QFF experiments is related to

the extensive side-lap region for this wing. With reference to figures 9a and 9b, the side-lap region is the

space near the flap side edge between the flap and the aft portion of the wing on the portion without a

flap. The side-lap region extends over approximately 40 percent of the flap chord.

Figure 10, which is taken from figure 6 of Berkman et al. (ref. 9), shows a cross-stream plane of

streamwise vorticity contours from a Reynolds averaged Navier-Stokes calculation of the flow. The view

in the figure is towards the upstream direction. The presence of the side lap produces a high-speed flat jet

of fluid that results in a strong shear layer that separates from the main element. This shear layer quickly

rolls up into a main-element vortex with a sign opposite to that of the flap-side-edge and flap-top vortices.

The presence of the main-element vortex appears to flatten the flap-top vortex and delay the merging of

the flap-side-edge vortex with the flap-top vortex relative to what was observed in the experiments

performed in the QFF. The configuration used in the QFF had a smaller side-lap region.

Figure5. Three-dimensionalviewofmainelementandflap.Dashedlineonflapshowsapproximatelocationoflineofpressuxetaps.

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Figure 6. Chordwise pressure distribution over wing model with single slotted flap. Spanwise location is

approximately 1 in. inboard of flap edge; c_ 5°; 8f 20°; Re 7.2 x 106; M 0.2.

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Figure 7. Pressure-sensitive paint measurements, c_ 5°; 8f 20°; Re 7.2 x 106; M 0.2.

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Figure 8. Vorticity contours from 5-hole pressure probe measurements in the QFF.

10

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(a) View from side.

Cut through trailing portionof main element

,----_ Side-lap region

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Cut through portion of flap1"lP" I

(b) View from downstream.Figure 9. Schematic of side-lap region between flap and aft portion of unflapped side of wing

As expected, all vortices are stronger with 8f = 30 ° than with 8f = 20 °. With 8f = 20 °, only very mildstrengthening of the vortices occurs with increases in the Reynolds number. The Reynolds number effect

appears to be slightly stronger, but still rather weak, with 8f = 30 °. Whether these changes with Reynoldsnumber are significant is not clear.

3.1.2. Unsteady Su_ace Pressures

The unsteady pressure transducer distribution used on the flap in the 1997 test is illustrated in

figure 11. In figure 11 the transducers are indicated by the rectangles with a darkened square at one end.

The darkened squares indicate the active sensing region of each transducer. The transducers on the flap

suction surface that are referenced later are labeled A-C. The two transducers on the flap side edge are

numbered 1 and 2. Autospectra from these transducers are shown in figure 12. These spectra show that

transducer 1 has two low-frequency peaks, suggesting that coherent oscillations of the flap-side-edge

vortex are present at its location. In contrast, the spectrum of transducer 2 is featureless, thereby

suggesting that the flap-side-edge vortex has already moved to the suction surface.

11

Vorticitynondimensionalized

5o25

---25

-50

Y

Z--X

-edge vortex

Figure 10. Color contours of streamwise vorticity from Reynolds averaged Navier-Stokes calculation, c_ 5°;

8f 20°; Re 7.2 x 106; M 0.2; view is looking upstream; plane is located at x/c 0.94; trailing portion of main

element is on left side; flap is on right side.

1 000 in _0 200 in

Figure 11. Schematic of unsteady pressure transducer distribution on flap side edge in LTPT tests. WRP is wing

reference plane. Transducers on the flap suction surface that are referenced later are labeled A-C. Transducers on

the flap side edge are labeled 1 and 2. Darkened squares are the active sensing regions.

12

10

Z"

0

o 10

E

2o

0 10 20 30 40 50

Frequency, kHz

Figure 12. Autospectra from transducers 1 and 2.8f 30°; Re 7.2 x 106; M 0.2.

Transducers A, B, and C are near the flap side edge on the suction surface (see fig. 11). As shown in

figure 13, the maximum amplitude signal is observed at transducer C. At this location, the flap-side-edgevortex has probably merged with the flap-top vortex. The spectrum of transducer C is the pressure

spectrum of the turbulence in the merged vortex system. Figure 14 shows the coherence betweentransducers B and C. The relatively strong coherence in the frequency range of 1500 to 4000 Hz suggests

that the same large scale flow structures are responsible for 10 to 30 percent of the pressure oscillations atboth locations.

2O

10

0

10o

2O

3O0 10 20 30 40 50

Frequency, kHz

Figure 13. Autospectra from transducers A, B, and C. 8f 30°; Re 7.2 x 106; M 0.2.

13

.4

.3

.2

i |

0 2.5 5.0 7.5

Frequency, kHz

Figure 14. Coherence between transducers B and C. 8f 30°; Re 7.2 x 106; M 0.2.

3.1.3. Acoustics

Acoustic measurements of noise originating at the flap side edge are consistent with previously

proposed ideas of instabilities in the turbulent shear layer at the side edge (refs. 21-23). The noise

generated at the flap side edge is most conveniently modeled as four sources that manifest themselves in

the frequency ranges illustrated schematically in figure 15. The integrated 1/12-octave sound pressure

level (SPL) on the vertical scale represents the 1/12-octave SPL integrated over a volume in the vicinity

of the flap side edge. The localization maps shown in figures 17-20 include a rectangular frame that

indicates the extent of a slice of the integration volume. The full integration volume includes similar

slices stacked in the direction normal to the slices shown. The acoustic sources are believed to be the

results of instabilities in the turbulent shear layers.

95-

9o

85

80

75

7o

650

Total -_

i

Vortex instability

-- Long wavelength cylindrical

shear-layer instabilities

U Short wavelength

l / shear-layer_-_ instabilities

/ _ /"x V Side-lap region

_i,,,s I

e

10 20 30 40 50 60

Frequency, kHz

Figure 15. Schematic integrated 1/12-octave acoustic spectra of turbulent flap-side-edge noise sources.

14

An actual spectrum of the SPL integrated over the flap-side-edge region is shown in figure 16. The

cutoff at 60 kHz is the result of a data management tradeoff and does not imply that no noise sources have

frequencies that exceed 60 kHz. The strongest noise source in figure 16 is associated with a strong

tone-like signal at slightly less than 5000 Hz.

Figures 17-20 show localization plots for run 61. The array centroid was approximately 1 m from the

flap suction surface. All localization plots shown here illustrate contours from the local maximum to

approximately 8 dB less than the maximum. Figure 17 shows a source localization plot of the noise at4870 Hz.

95-

9o

85

80

75

7o

65 i10 20 30 40 50 60

Frequency, kHz

Figure 16. Integrated 1/12-octave acoustic spectrum for run 61 of 1998 test. 8f 30°; Re 7.2 x 106; M 0.2.

Figure 17. Localization plot for run 61 of 1998 test. f 4870 Hz; 8f 30°; Re 7.2 x 106; M 0.2; local1/12-octave SPL peak is 92.5 dB.

15

Figure18. Localizationplotfor run61of 1998test.f 9173 Hz; 8f 30°; Re 7.2 x 106; M 0.2; local1/12-octave SPL peak is 70.0 dB.

IFigure 19. Localization plot for run 61 of 1998 test. f 16312 Hz; 8f 30°; Re 7.2 x 106; M 0.2; local1/12-octave SPL peak is 62.8 dB.

The source localization plot in figure 17 shows that a strong source exists at the trailing corner of the

flap side edge. This noise source is believed to be caused by interaction of unsteady flow structures in the

merged flap-side-edge vortex with the flap trailing edge. The calculations of Streett (ref. 22) show the

development of coherent, ringlike flow structures in the outer portion of the merged vortex. This noise

source was effectively eliminated by the introduction of the noise weeder described below.

16

Figure 20. Localization plot for run 61 of 1998 test. View from pressure side ofmodel.f 36517 Hz; 8f 30°;Re 7.2 x 106; M 0.2; local 1/12-octave SPL peak is 54.9 dB.

As frequency increases, the dominant noise source tends to move upstream on the flap. Figures 18-20

show source localization maps at frequencies of 9173 Hz, 16312 Hz, and 36517 Hz, respectively. In the

figures, the plane on which the contours are displayed intersects the flap suction surface. Because the

array has poor resolution in the direction normal to the wall on which it is mounted, noise sources might

appear to emanate from inside the solid model. However, in figure 18, no data are illustrated in the

intersection region of the solid model and the plane of the localization map.

At all conditions, some noise radiates from the trailing-edge corner, but another important source

appears on the side edge. The regions associated with flap-side-edge peaks in figures 18 and 19 are

believed to result from instabilities in the turbulent shear layer at the flap side edge. These instabilities are

consistent with the stability analyses of Khorrami and Singer (refs. 21 and 24) and the detailed numerical

simulations of Streett (ref. 22). The calculations of Streett (ref. 22) clearly show the different nature of the

instabilities in the high- and low-frequency ranges.

The source of the highest frequency noise at the flap side edge appears to be a vortex roll-up and the

associated turbulent flow. This phenomenon results from the jet-like flow through the side-lap region

between the pressure surface of the main element's trailing edge and the suction surface of the flap's

leading edge. (Refer to figure 9 for the geometry.) The localization plot in figure 20 is shown from a

slightly different angle than that of figures 17-19 to better reveal that the upstream noise source emanates

from the side-lap region. This noise source is likely to be important only in high-lift devices with

significant amounts of side lap.

3.1. 4. Flap-Side-Edge Modifications

During the 1997 test, a variety of flap-side-edge modifications were tried, which primarily involved

altering the shape of the flap side edge. Figure 21 shows the different flap edges used. The baseline flap

edge (first from left in fig. 21) was flat, resulting in a sharp comer at both the lower and upper edges.

17

Figure21.Differentflapsideedgesusedin1997tests,fromlefttoright:baselineflapedge,half-round1,flangefilledwithporousmaterial,full-round,half-round2,andhalf-round3.

Threeflapsideedgeswithsharpcornersattheupperedge,butwithroundedloweredge,aredenotedhalf-round1,half-round2, andhalf-round3. Thefull-roundwasdesignedwiththeentireflapsideedgerounded,sothatneitherthetopnorbottomcornerswassharp.Theflangemodificationwasaflatflapside-edgewitharecessedcenterportion(notshown).A finalmodification"shown in fig. 21) involved the use

of the flange with the recessed portion filled with a porous liner.

At low Reynolds numbers, the half-round configurations were noisier than the baseline, while the full-

round configuration was slightly quieter than the baseline. With Re = 14.4 x 106, where all of the

modifications were tested, the half-round 1 modification remained somewhat noisier than the baseline,

while the other modifications appeared to make no substantial change in acoustic radiation compared with

the baseline edge. Figure 22 shows the spectra of the half-round modifications plotted together with the

baseline case. Figure 23 shows the spectra for the remaining edge modifications and the baseline case.

The small gaps in the data at about 4590 Hz and 23040 Hz are associated with the data processing

algorithm used at that time.

With the knowledge gained from the 1997 test, two concepts for flap-edge modifications were tested

in 1998. Based on the hypothesis that much of the noise at the side edge was associated with instabilities

in the turbulent shear layer, a technique was developed to modify the shear layer to reduce the growth of

these instabilities. The technique employed microtabs positioned near the flap side edge on the flap

pressure surface. The microtabs are small trapezoidal devices that shed vortices from their corners and

thereby increase the thickness of the resulting detached shear layer. A strip of microtabs attached to the

pressure surface of the flap side edge is shown in figure 24. Detailed dimensions of typical microtabs

18

sizedforthe1998flap-side-edgenoisereductiontestsareshownin figure25.105

100

95

90

85

8O

75

70

65

6O

0

-- Baseline

.... Hall'-rowed 1-- - -- Hall'-rowed 2

.......... Hall'-rowed 3

10 20

Frequency, kHz

30 504O

Figure 22. Integrated 1/12-octave acoustic spectra from 1998 test. 8f 30°; Re 14.4 x 106;M 0.2.

105

100

95

9O

85

80

75

7O

65

60

-- Baseline

.... Full-round

---- Flange.......... Flange with liner

i i i I i i

10 20 30 40 50

Frequency, kHz


19

Figure 24. Microbtabs on flap side edge.

W

,._------ a _________ b_

Figure 25. Schematic of typical microtab strip on flap side edge. For the LTPT experiments, a 0.075 in.,b 0.075 in.,c 0.051 in.,d 0.060in.,e 0.012in.,h 0.046in.,L 4in., W 0.10in.,c_ 50 °.

The microtabs used in this application thicken the mixing region and decrease mean shear. The shear

layers are therefore more stable because microtabs produce more mixing as a result of greater entrainment

associated with small vortices generated by the microtabs.

Previous data indicate that a significant portion of the noise associated with the flap side edge radiates

from the trailing-edge corner of the flap. The microtabs were designed to thicken the separated shear

layer. Figure 26 shows results with and without the microtabs for runs with 8j. = 20 °. Noise wassignificantly reduced from approximately 7 kHz to 40 kHz. However, the microtabs apparently increase

the noise between 40 and 50 kHz, probably by adding small-scale fluctuations to the flow. Details

associated with microtab sizing and placement need to be explored more carefully to make this

modification viable. Tests with the same placement of the microtabs and 8j. = 30 ° showed essentially no

20

90

85

8O

75

-- Baseline (run 49) V " \..... Flap microtabs in use (rim 20)

60 i i i i I i i i i I i llllllllhmhml

0 10 20 30 40 50 60

Frequency, kHz

Figure 26. Integrated 1/12-octave acoustic spectra from 1998 test. 8f 20°; Re 7.2 x 106; M 0.2.

noise reduction. For 8j. = 30 °, appreciable noise reduction was achieved with the microtabs placed fartherinboard of the flap side edge.

More recently, relevant experiments were performed as follows: (1) a trapezoidal wing equipped with

a high-lift system in the Langley 14- by 22-Foot Subsonic Tunnel (test 480) and (2) a separate 2D wing

with a part-span flap in the Langley Basic Aerodynamic Research Tunnel (test 52). Results of these

experiments suggest that maximum noise reduction is achieved with flap-side-edge microtabs placed two

to four microtab heights inboard of the flap side edge.

Extensive evidence indicates that much of the noise associated with the flap side edge radiates from

the vicinity of the trailing edge. This phenomenon is most clearly evident in the case of the low-frequency

tone shown in the spectrum of figure 16 and in the localization map in figure 17. Eliminating, or at least

reducing, the level of this tone is critical to obtaining any substantial noise reduction.

Serrations in wing trailing edges have been proposed by Howe as a noise-reduction technique

(ref. 25). In reference 25, the proposed serrations covered the full extent of the wing's trailing edge, and

the work did not investigate whether aerodynamic performance had been degraded by the use of the

serrations. Flaps are part of the aircraft's high-lift system and are subject to detailed scrutiny during the

flight certification process. Any noise-reduction modification must guarantee that it induces negligible

degradation to the system's aerodynamic performance.

A proposed modification involved the inclusion of a short span of trailing-edge serrations near the flap

side edge. Force-balance data showed less than 0.5 percent change in lift with all of the proposed noise

reduction modifications. The modified side edge spanned 2 in. Figure 27 shows a photograph of the

device, which we call a "noise weeder" because of an appearance similar to certain garden implements.

The specific noise weeder that we tested included three full serrations and a half-tooth terminating each

spanwise end of the modified section. Each full serration spanned 0.5 in., cut 0.5 in. into the unmodified

edge, and extended 0.5 in. beyond the unmodified edge. With these dimensions, the sweep angle of the

serration tip was slightly more than 60 ° and the serrations maintained the same surface area as the

21

Figure27.Noiseweederonflapsideedge.

unmodifiededge,i.e.,regionswheretheserrationscutintothetrailingedgewerebalancedbyregionswheretheserrationsextendedbeyondtheunmodifiedtrailingedge.

This"equivalentsurfacearea"approachis believedto be importantto maintainingaerodynamicperformance.Furthertestswouldbenecessaryto determinethedegreeto whichdeviationsfromtheequivalentsurfaceareaguidelinecouldbe tolerated.Becausetheunmodifiedflap includeda finite-thicknesstrailingedge,projectingthegeneralcontouroftheairfoilprofileintotheextendedportionof theserrationwasstraightforward.

Theserrationdimensionswerechosento minimizethespanoverwhichserrationswerenecessarytoachieveflap-side-edgenoisereduction.The2-in.spanof themodificationwaschosento includethepathof theflap-side-edgevortexas it rolledto thesuctionsurfacenearthetrailingedge.Thisspanwasinferredfromprevioustests(forinstance,seefigure6) andcalculations(ref.12).By assumingthattheflow featuresscalewiththegeometry,webelievethatanappropriatespanfor afull-scaleflap,wouldbetwicethedistancefromthecenterlineoftheflap-side-edgevortextotheflapsideedgeatthepointwheretheside-edgevortexpassestheflaptrailingedge.

Althoughfurtherexperimentationmightrevealothersuccessfuloptions,the"threefull teeth,twohalf-tooth"configurationisbelievedtobeagoodchoiceevenforafull-scaleflap.Theuseof thehalf-toothateachspanwiseendof themodifiedregionfacilitatesmatingthemodifiedregionwith theunmodifiedregion.Thehalf-toothalsoprovidesfor a cleanflapsideedge.Variationsto allowfor wingsweepandotherthree-dimensionaleffectscanbeincorporatedaswell.Theapproximate60° sweep angle of each

serrated tip is believed to be a good choice that balances the acoustic scattering reduction effect achieved

with high sweep and the practical problems associated with having numerous narrow teeth.

22

Figure 28 compares spectra with and without the noise weeder for the case of 8j. = 20 °. The low-frequency peaks are effectively removed by the noise weeder. A slight increase in noise above about16 kHz is indicated.

9O

85

8O

75

7O

65

60

0

;' ...... 'A/v- Baseline (run 49) V v

,. Noise)_'eeder(,run,37,), ,,, .... , .... ,,,,,,,,,,,

10 20 30 40 50 60

Frequency, kHz

Figure 28. Integrated 1/12-octave acoustic spectra from 1998 test. 8f 20°; Re 7.2 x 106; M 0.2.

Because of time restrictions during the 1998 test, the noise weeder modification alone was not tested

with 8j. = 30 °. Instead, the noise weeder was tested in combination with microtabs. Figure 29 shows theintegrated acoustic spectra for a baseline case and a case with the noise weeder and microtabs. In this

case, microtabs were included on the main element pressure surface in the side-lap region (see fig. 30)

and on the noise weeder pressure surface slightly inboard of the side edge (see fig. 31). The noise weeder

95-

A9O

[/ Baseline (run 61)

85 I I Noise weeder

} / plus microtabs (run 93)

/ \_

,,' ' ,r,/65 I_ .... I .... I .... I .... h,,,h&

10 20 30 40 50 60

Frequency, kHz

80


23

Figure 30. Microtabs on pressure surface of main element in side-lap region. View is from trailing edge of flap

towards side-lap region.

Figure 31. Microtabs on side edge of noise weeder.

24

essentially eliminated the low-frequency tone that existed at just under 5 kHz. In addition, some

improvement in the noise was achieved in the mid-frequency range where the shear layer instabilities are

important. With regard to the high-frequency noise, the data show essentially no change in the noise.

Further optimization of the noise-reduction techniques could be expected to provide additional reductions.

Figures 32-36 show localization plots for run 93. Along with the noise weeder, this run used

microtabs on the main element pressure surface and on the pressure surface of the flap side edge. As for

the baseline case in figures 17-20, the array centroid was approximately 1 m from the flap suction

surface. Except for the additional frequency of 23 041 Hz in figure 35, the frequencies are the same as

those shown for the baseline case. Although the actual levels of the contours differ in each figure, the

decibel range is approximately the same in all the localization plots. Comparison between the baselinecase and the modified case shows how the modifications have altered the locations of the dominant noise

sources.

Comparing figures 17 and 32 shows that at 4870 Hz, the modifications removed the maximum that

was slightly downstream of the flap trailing edge and left a maximum slightly upstream of the flap trailing

edge. A significant reduction in the maximum SPL also occurred. Spectra shown in figure 28 indicate that

the presence of the noise weeder alone removed the low-frequency tone. At 9173 Hz, figure 33 shows the

maximum at the flap side edge from figure 18 was reduced, leaving two maxima, one slightly upstream

and one slightly downstream. A reduction in the maximum also occurred for 16312 Hz. This reduction

was accompanied by the elimination of the most upstream peak in figure 19; this peak does not appear in

figure 34. A localization plot at 23041 Hz (fig. 35) is included for the case with the edge modifications

because the noise at this frequency was greater with the modifications than without them. The dominant

noise source appeared to be at the flap side edge at about 2/3 chord. Some noise that might be considered

side-lap noise developed further upstream. A minor noise source appeared slightly downstream of the

trailing edge. At 36517 Hz, the localization plot in figure 36 shows a considerable reduction in side-lap

noise compared to the unmodified flap side edge shown in figure 20.

Figure 32. Localization plot for run 93 of 1998 test. f 4870 Hz; 8f 30°; Re 7.2 x 106; M 0.2. Local1/12-octave SPL peak is 71.2 dB.

25

Figure 33. Localization plot for run 93 of 1998 test. f 9173 Hz; 8f 30°; Re 7.2 x 106;M 0.2; local

1/12-octave SPL peak is 67.3 dB.

Figure 34. Localization plot for run 93 of 1998 test. f 16312 Hz; 8f 30°; Re 7.2 x 106; M 0.2; local


26

Figure 35. Localization plot for run 93 of 1998 test. f 23 041 Hz; 8f 30°; Re 7.2 x 106; M 0.2; local


Figure 36. Localization plot for run 93 of 1998 test. f 36517 Hz; 8f 30°; Re 7.2 x 106;M 0.2; local


27

These observations suggest that the noise weeder removed the low-frequency tone associated with the

flap-side-edge vortex traversing the flap trailing edge. In addition, the microtabs on the pressure surface

of the main element appeared to have significantly reduced the side-lap noise, and the microtabs on the

flap side edge appeared to have reduced the shear layer noise. The noise weeder might have introduced

some high-frequency noise (the small source downstream of the trailing edge), as was shown in the

8f = 20 ° spectra of figure 28, but the 8f = 30 ° spectra in figure 29 show little change in the highfrequencies.

3.2. Leading-Edge Slat

3.2.1. Mean Flow

A typical static pressure distribution on the slat is shown in figure 37. For this case, the slat was

deflected with 8 s = 30 °. Flow on the pressure side of the slat separated at the slat cusp, but reattached

upstream of the trailing edge. Other experiments on a similar configuration found the reattachment was

unsteady and could produce strong fluctuations in the flow field (ref. 26).

Much of the mean flow information about the slat flow field was derived from computational studies

(refs. 9 and 10). The computations were three-dimensional (3D) and included the part-span flap, but not

the flap nor slat brackets. Figure 38 shows streamlines superimposed on Mach contours in the region

around the slat. The large recirculation region is evident in the figure. A free-shear layer developed on the

edge of the recirculation zone. Fluid was accelerated through the gap at maximum speeds almost 2.5

times the free-stream velocity. Another important observation was that the slat flow field was essentially

2D, in spite of the inclusion of the part-span trailing flap. However, the presence of the slat brackets in the

wind tunnel experiments most certainly introduced at least local 3D effects into the flow.

5

4

3

1

0

1.1

illl

.1x/7:

Figure 37. Pressure coefficient measured on slat centerline, a 10°; 8f 30°; 8s 30 °.

28

Figure 38. Streamlines superimposed on Mach contours in vicinity of slat. Spanwise location is 1 in. inboard from

flap side edge. cz 10°; 8f 30°;8 s 30°;M 0.2.

3.2.2. Unsteady Su_ace Pressures

The locations of five unsteady pressure transducers used on the slat in the 1998 test are illustrated in

figure 39. Figure 40 shows the spectra from transducers 1 and 5 and from a single microphone of the

microphone array. The broad peak above 40 kHz in the microphone signal is probably related to the

unsteady vortex shedding at the slat trailing edge, as discussed below. The surface pressure signal also

appears to increase slightly in this frequency range.

J

/Transducer 5

Transducer 1

Figure 39. Schematic of five unsteady pressure transducer locations on slat.

29

140

%-.

o

130

120

110

IO0

9O

8O

7O

600

Microphone

20 40 60 80

Frequency, kHz

Figure 40. Signals from transducers 1 and 5 on slat, plus a microphone of the acoustic array.

3.2.3. Acoustics

During the 1998 tests, the slat noise spectra were dominated by a high-amplitude, high-frequency

peak. Figure 41 shows the spectra for two slat deflections. The high-frequency peak is very clear for8s = 30°, but is substantially reduced for the 8s = 20° case. Khorrami et al. (ref. 10) hypothesized that

vortex shedding at the slat trailing edge was responsible for the high-frequency noise. Their 2D

- ;_0

m 85

80-

75

7o10 20 30 40 506070

Frequency, kHz

Figure 41. Integrated 1/12-octave spectra on slat. c_ 10°; Re 7.2 x 106;M 0.2.

30

unsteady calculations support this theory. With this unsteady data as input, the acoustic analogy

calculations of Singer, Lockard, and Brentner (ref. 8) demonstrated that the vortex shedding produced

sufficient noise to explain the peak and that the directivity of the computed acoustic signal was consistent

with individual microphone data from the experiment.

Storms et al. (ref. 27) noted a similar strong sensitivity to slat deflection in experiments on a different

high-lift system. They attributed the difference to a laminar turbulent transition on the suction surface of

the slat. At the lower slat deflections, a laminar separation bubble had a turbulent reattachment, while for

the higher slat deflections, no separation bubble was expected and the flow remained laminar. On the

suction surface, a laminar boundary layer would have been more prone than a turbulent one to vortex

shedding at the trailing edge and hence could account for the production of the high-frequency tone at the

higher slat deflections. However, the current study was conducted at a variety of significantly higher

Reynolds numbers. Figure 42 shows the spectra for four different Reynolds numbers with M = 0.2 and

8 s = 30 °. The loud tone shows some variability for the two lower Reynolds numbers, but is remarkably

similar for the two higher Reynolds numbers. Therefore, a transition effect would not likely account forthe difference between the two slat deflections.

Because the geometry of the 20 ° slat deflection differed from that of the 30 ° slat deflection, a

resonance mechanism that might be involved in the slat tone noise for the 30 ° slat deflection probably

would not be applicable to the case with the 20 ° slat deflection. Such resonances were considered as

potentially amplifying the noise generated by the vortex shedding. No simple resonance theory has yet

been developed.

The 1999 tests were designed to confirm the source of the high-frequency peak and to explore a

method for reducing the low-frequency noise. To aid in the investigations, an additional small-aperture

95

9O

85

8O

75

7O

Re /

3.6 × 106 if\ ;7.2 × 10% _ '-"d

14.4x 10 0 _'_ _1'

........... 19°×1°° i_'/ l

i i i i _""1 i i i _\_ i i_/i llllllllllll ._d

10 20 30 40 506070

Frequency, kHz

Figure 42. Integrated 1/12-octave spectra on slat. c_ 10°; 8s 30°;M 0.2.

31

arraywasusedin conjunctionwith thelarge-aperturearray.Thesmall-aperturearraywasdesignedtofocusmoredistinctlyonhigh-frequencynoiseemanatingfromthevicinityof theslat.Thearraycentroidwasapproximately1m fromthepressuresurface,slightlyupstreamof theslatleadingedgeandbiasedoff thewindtunnelcenterline.An acousticlocalizationmapisshownin figure43.In contrast,thelarge-aperturearrayhaditscentroidslightlyupstreamofthequarterchordportionofthemainelement.

Althoughthesmall-aperturearraywasdesignedprimarilyto collecthigh-frequencydata,for manycasessmall-aperturearraydatawerecollectedinsufficientlylongsegmentsforthedatatobeprocessedtolowerfrequencies.Baselineandmodifiedwingdataareshownonlyoverfrequencyrangesinwhichdatawereavailableforbothcases.

3.2.4. Slat Modifications

Three modifications were made to the basic slat configuration. Two of the modifications were

designed to reduce the noise emanating from the slat trailing edge; the third was designed to reducebroadband cove noise.

The first modification involved the placement of serrated tape near the trailing edge of the suction

surface of the slat. Figure 44 shows the tape on the slat. Figure 45 shows a close-up view that indicates

roughly the dimensions involved. The intended purpose of the tape was to thicken the boundary layer on

the slat and thereby eliminate or reduce the vortex shedding that was so dominant on the baseline model.

A comparison of the microphone array spectra, with and without the tape, is shown in figure 46. The use

of the serrated tape reduced the high-frequency peak by approximately 6 to 7 dB.

The second slat modification replaced the trailing edge of the slat with a sharper trailing edge. Rather

than having a trailing-edge thickness of approximately 0.020 in., the nominal trailing-edge thickness was

reduced to approximately 0.010 in. Unfortunately, this slat trailing edge was not uniform across the span.

dB from peak

N0:.:.:::::::.---2..::= 4

Figure 43. Localization plot from small-aperture acoustic array for run 38 of 1999 test. f 48697 Hz; c_ 6°;

8s 30°; Re 7.2 x 106; M 0.2; local 1/12-octave SPL peak is 80.3 dB.

32

Figure 44. Serrated tape on suction surface of slat. Because model is inverted, view is from floor looking towards

tunnel ceiling.

Figure 45. Close-up view of serrated tape on suction surface of slat.

33

9O

85

8O

75

7O

65

60

-- Baseline large-aperture arrayLarge-aperture array wi/h serrated tape

.... Baseline small-aperture array

- -El- - Small-aperture array with serrated tape

z \

/ \

/

1

I _.\ _

/ ., \',,,

";"S

........57 ....LT10 20 30 40 50 60 70

Frequency, kHz

Figure 46. Integrated 1/12-octave spectra on slat. c_ 6°; 8s 30°; Re 7.2 x 106;M 0.2.

A tone at approximately 14 kHz developed at a number of conditions with the thinner slat trailing edge.

Although the exact cause of the tone is not yet understood, figure 47 is a localization plot at 14 kHz that

suggests the source of this tone is near center-span of the slat. Therefore, this tone does not appear to be

associated with bracket noise, but may be associated with the nonuniform trailing-edge thickness.

Figure 47. Localization plot from large-aperture acoustic array for run 98 of 1999 test; thin trailing edge;

f 14538 Hz; c_ 6°; 8s 30°; Re 7.2 x 106;M 0.2; local 1/12-octave SPLpeakis 81.1 dB.

34

However, the tone does not appear for cases with Re = 14.4 x 106. Although a deeper understanding of

the 14 kHz tone is desirable, the higher Reynolds number more closely approximates the Reynolds

number range of a commercial aircraft on approach. Figure 48 shows that at this higher Reynolds number

the noise is greatly reduced from the baseline case. The reduction is almost 10 dB in the frequency range

associated with the vortex shedding noise.

The above results indicate that certain slat modifications dramatically reduced the noise produced in

the vicinity of the 50 kHz tone that was observed in the baseline case. The slat modifications that resulted

in such reductions were designed to either increase the slat-trailing-edge boundary-layer thickness or

decrease the slat-trailing-edge thickness. This experimental observation strongly supports the

computational conclusions in references 8 and 10 that vortex shedding was responsible for the 50 kHztone.

The third slat modification was designed to reduce the lower frequency broadband noise. This

modification was similar to the partial slat coverings used by Dobrzynski et al. (ref. 3) and shown in

reference 28. However, instead of a partial slat covering, a closed surface filler was inserted into the slat

cove. Previous experiments that employed slat-cove fillers had used the separation streamline to define

the filler geometry (refs. 29 and 30). However, the use of such a filler geometry resulted in the formation

of a large separation bubble on the aft portion of the filler and the generation of a significant amount of

high-frequency noise.

In the current study, the design philosophy was to fill the cove region in such a way that the gap

between the slat and the main element was a smoothly converging duct. A cross section of the slat and the

filler is shown in figure 49.

-- Base]ine large-aperture array90 ------Ell Large-aperture array with thin slat lrai]ing edge

_- .... Base]ine small-aperture arrayL - --D- - Sma]l-aperture array with thin slat lrai]ing edge

i1_

85. / _

80 _

_c__75 \\

70 t -- ' //A _ x]_ z, M_ ,'-_'-'t_

6O0 10 20 30 40 50 60 70

Frequency, kHz

Figure 48. Integrated 1/12-octave spectra on slat. a 6°; 8 s 30°; Re 14.4 x 106;M 0.2.

35

I

i I

Figure 49. Schematic of slat (solid line) with slat-cove filler (dashed line).

The slat-cove filler was used only with the thinner slat trailing edge. This configuration was expected

to be the quietest. The spectra obtained are compared with a baseline case in figure 50. The noise

reduction achieved with the slat filler was highly dependent on frequency. A reduction of approximately

10 dB was observed over a significant portion of the 0-20 kHz range.

75

c_

-- Baseline large-aperture arrayLarge-aperttu'e array with thin slat

90 _ trailing edge and slat-cove filler- - - Baseline small-apertt_re array

- '[3- - Sma|l-aperture arrav with thin slattrailing edge and slat-cove filler t_

70 _\

65-

60 _ I_1_1_1_1_1_1

0 10 20 30 40 50 60 70

Frequency, kHz

Figure 50. Integrated 1/12-octave spectra on slat. c_ 6°; 8 s 30°; Re 14.4 x 106;M 0.2.

36

4. Concluding Remarks

This report describes details of experimental airframe noise investigations performed in the Langley Low-

Turbulence Pressure Tunnel. Reynolds numbers ranging from 3.6 x 10 6 to 19.2 x 10 6 were explored. In

the various tests, steady and unsteady surface pressure data were obtained in addition to phased

microphone array data. This work focused on noise emanating from the flap and slat portions of a

high-lift wing.

The flap side edge produces a complicated flow comprising multiple vortices that propagate

downstream and merge. When the data were evaluated in conjunction with associated numerical

simulations, four primary mechanisms were suggested for the generation of noise from the flap side edge.

Flap-side-edge modifications that interfered with these mechanisms were shown to reduce the radiated

noise without incurring significant aerodynamic penalties.

Noise radiation from the slat still is not well characterized. One mechanism is associated with vortex

shedding from the slat trailing edge. Other mechanisms probably are associated with the separated flow in

the slat cove, but these mechanisms are not yet clearly understood. Modifications to the slat trailing edge

can essentially eliminate the noise from the vortex shedding. A slat-cove filler significantly reduced the

low-frequency noise by essentially eliminating the separation zone in the slat cove.

37

References

1. Davy, H.; and Remy, R.: Airframe Noise Characteristics on a 1/11 Scale Airbus Model. AIAA-98-2335, 1998.

2. Hayes, J. A.; Horne, W. C.; Soderman, P. T.; and Bent, P. H.: Airframe Noise Characteristics of a 4.7% Scale

DC-10 Model. AIAA-97-1594, 1997.

Dobrzynski, W.; Nagakura, K.; Gehlhar, B.; and Buschbaum, A.: Airframe Noise Studies on Wings With

Deployed High-Lift Devices. AIAA-98-2337, 1998.

Macaraeg, M. G.: Fundamental Investigations of Airframe Noise. AIAA-98-2224, 1998.

Meadows, K. R.; Brooks, T. F.; Gerhold, C. H.; Humphreys, W. M. Jr.; and Hunter, W. W.: Acoustic and

Unsteady Surface Pressure Measurements of a Main Element-Flap Configuration. AIAA-97-1595, 1997.

6. Radeztsky, R. H.; Singer, B. A.; and Khorrami, M. R.: Detailed Measurements of a Flap Side-Edge Flow Field.

AIAA-98-700, 1998.

7. Brooks, T. F.; and Humphreys, W. M. Jr.: Flap Edge Aeroacoustic Measurments and Predictions. AIAA-2000-

1975, 2000.

8. Singer, B. A.; Lockard, D. P.; and Brentner, K. S.: Computational Acoustic Analysis of Slat Trailing-Edge

Flow. AIAA J., vol. 38, no. 9, Sept. 2000, pp. 1558 1564.

9. Berkman, M. E.; Khorrami, M. R.; Choudhari, M.; and Sadowski, S. S.: Investigation of High-Lift Flow Field

of an Energy Efficient Transport Wing. AIAA-99-0926, 1999.

10. Khorrami, M. R.; Berkman, M. E.; Choudhari, M.; Singer, B. A.; Lockard, D. P.; and Brentner, K. S.: Unsteady

Flow Computations of a Slat With a Blunt Trailing Edge. AIAA-99-1805, 1999.

11. McGhee, R. J.; Beasley, W. D.; and Foster, J. M.: Recent Modijications and Calibration of the Langley Low-

Turbulence-Pressure Tunnel. NASA TP-2328, 1984.

12. Morgan, H. L.: Model Geometry Description and Pressure Distribution Data From Tests' of EET High-L_

Research Model Equipped With Full-Span Slat and Part-Span Flaps. NASA TM-80048, 1979.

13. Lin, J. C.; and Dominik, C. J.: Parametric Investigation of High-Lift Airfoil at High Reynolds Numbers. J.

Aircr., vol. 33, no. 4, 1997, pp. 485 491.

14. Underbrink, J. R.; and Dougherty, R. P.: Array Design for Non-Intrusive Measurements of Noise Sources.

Proceedings of NOISECON 96, J. D. Chalupnik, S. E. Marshall, and R. C. Klein, eds., Institute of Noise Control

Engineering, Washington, DC, 1996.

15. Dougherty, R. P.: Spiral-ShapedArrayJorBroadbandImaging. U.S. Patent No. 5,838,284, 1998.

16. Mosher, M.: Phased Arrays for Aeroacoustic Testing_heoretical Development. AIAA-96-1713, 1996.

17. Mosher, M.; Watts, M. E.; Jaeger, S. M.; and Jovic, S.: Calibration of Microphone Arrays for Phased Array

Processing. AIAA-97-1678, 1997.

18. Radeztsky, R. H.; Singer, B.S.; and Khorrami, M. R.: Detailed Measurements of a Flap Side-Edge Flow Field.

AIAA-98-0700, 1998.

3.

4.

5.

38

19. Khorrami, M. R.; Singer, B. A.; and Radeztsky, J.: Reynolds-Averaged Navier-Stokes Computations of

Flap-Side-Edge Flowfield. AIAA J., vol. 37, no. 1, 1999, pp. 14 22.

20. Storms, B. L.; Takahashi, T. T.; and Ross, J. C.: Aerodynamic Influence of a Finite-Span Flap on a Simple

Wing. SAE Paper No. 951977, 1995.

21. Khorrami, M. R.; and Singer, B. A.: Stability Analysis for Noise-Source Modeling of a Part-Span Flap. AIAA-

98-2225, 1998.

22. Streett, C. L.: Numerical Simulation of Fluctuations Leading to Noise in a Flap-Edge Flowfield. AIAA-98-

0628, 1998.

23. Streett, C. L.: Numerical Simulations of a Flap-Edge Flowfield. AIAA-98-2226, 1998.

24. Khorrami, M. R.; and Singer, B.A.: Stability Analysis for Noise-Source Modeling of a Part-Span Flap. AIAA

J., vol. 37, no. 10, 1999, p. 1206.

25. Howe, M. S.: Aerodynamic Noise of a Serrated Trailing Edge. J. Fluids' & Struct., vol. 5, 1991, pp. 33 45.

26. McGinley, C. B.; Anders, J. B.; and Spaid, F. W.: Measurements of Reynolds Stress Profiles on a High-Lift

Airfoil. AIAA-98-2620, 1998.

27. Storms, B.; Hayes, J.; Moriarty, P. J.; and Ross, J.: Aeroacoustic Measurements of Slat Noise on the Three-

Dimensional High-Lift System. AIAA-99-1957, 1999.

28. Heller, H. H.: DLR's Involvement in European Aviation Noise Research on Fixed and Rotary Wing

Aircraft A (Roughly) Five Year Retrospective. 5 th AIAA/CEAS Aeroacoustics Conference Keynote Lecture,

AIAA, Reston, VA, 1999.

29. Moriarty, P.: Unsteady Measurements Near a Leading-Edge Slat. 3 rJ AirJ?ame Noise Workshop, R. Sen, ed.,

NASA AST Program, 1998.

30. Storms, B.; Hayes, J.; and Ross, J.: Aeroacoustic Measurements of Slat Noise on a Three-Dimensional High-

Lift System. 3 rJ AirJ?ame Noise Workshop, R. Sen, ed., NASA AST Program, 1998.

39

Form ApprovedREPORT DOCUMENTATION PAGE OMSNo.0704-0188

Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources,gathering and maintaining the data needed, and completing and reviewing the collection of information. Send comments regarding this burden estimate or any other aspect of thiscollection of information, including suggestions for reducing this burden, to Washington Headquarters Services, Directorate for Information Operations and Reports, 1218 JeffersonDavis Highway, Suite 1204, Arlington, VA 22202-4302, and to the Office of Management and Budget, Paperwork Reduction Project (0704-0188), Washington, DC 20803.

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February 2002 Technical Memorandum4. TITLE AND SUBTITLE 5. FUNDING NUMBERS

Aeroacoustic Experiments in the NASA Langley Low-Turbulence PressureTunnel WU 706-81-13-02

6. AUTHOR(S)

Meelan M. Choudhari, David R Lockard, Michele G. Macaraeg, Bart A.

Singer, Craig L. Streett, Guy R. Neubert, Robert W. Stoker, James R.

Underbrink, Mert E. Berkman, Mehdi R. Khorrami, and Shelly S. Sadowski

7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)

NASA Langley Research CenterHampton, VA 23681-2199

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES)

National Aeronautics and Space AdministrationWashington, DC 20546-0001

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REPORT NUMBER

L-18131

10. SPONSORING/MONITORING

AGENCY REPORT NUMBER

NASA/TM-2002-211432

11. SUPPLEMENTARY NOTES

Choudhari, Lockard, Macaraeg, Singer, and Streett: Langley Research Center, Hampton, VA; Neubert, Stoker, andUnderbrink: Boeing Commercial Aircraft, Seattle, WA; Berkman, Khorrami, and Sadowski: High TechnologyCorporation, Hampton, VA.

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Unclassifie_UnlimitedSubject Category 71 Distribution: NonstandardAvailability: NASA CASI (301) 621-0390

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13. ABSTRACT (Maximum 200 words)

A phased microphone array was used in the NASA Langley Low-Turbulence Pressure Tunnel to obtain acousticdata radiating from high-lift wing configurations. The data included noise localization plots and acoustic spectra.The tests were performed at Reynolds numbers based on the cruise-wing chord, ranging from 3.6 x 106 to19.2 x 106. The effects of Reynolds number were small and monotonic for Reynolds numbers above 7.2 x 106.

14. SUBJECTTERMS

Aeroacoustic noise; Flap noise; Slat noise; Flap-edge noise; Flap-side-edge noise;Airframe noise; High lift; Acoustic array

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